It is known that the efficiency of a combustion turbine engine will improve as the firing temperature of the combustion gas is increased. The firing temperature of modern gas turbine engines exceeds the safe operating temperature of the metals used to form many of the engine components. As a result, it is known to apply a thermal barrier coating to protect an underlying metal substrate in portions of the engine exposed to extremely high temperatures.
Ceramic thermal barrier coatings are typically applied using a high temperature process, such as APS or EB-PVD. In a typical application, a nickel or cobalt based super alloy substrate is first coated with a layer of a bond coat material, such as MCrAlY or alumina, and then insulated with a layer of ceramic material. Typical ceramic insulating materials include zirconium oxide or hafnium oxide stabilized with yttrium. The thermal barrier coating is attached to the underlying bond coat and substrate along a single interface plane. Such coatings, however, are prone to failure due to the cracking and eventual separation or spalling of the ceramic coating from the underlying substrate caused by differing coefficients of thermal expansion and other causes. Accordingly, this leads to a conservative engine design because the integrity of the thermal barrier coating cannot be assured. Such conservative designs tend to utilize more cooling air and less thermal barrier coatings to lower the firing temperature, or otherwise lower the overall engine efficiency.
U.S. Pat. No. 4,639,388 dated Jan. 27, 1987, describes a method for reinforcing a thermal barrier coating so that the ceramic material is secured to the substrate along more than the single interface plane. That patent describes a metallic honeycomb structure that is metallurgically bonded to the substrate, with ceramic insulating material then deposited within the cells of the honeycomb structure. In addition to a honeycomb structure, other forms of metallic reinforcing structures are described in the patent, including pins, tabs, and coiled wires. The reinforcing structure may be formed to be integral with the substrate, or it is brazed, welded or diffusion bonded to the substrate material. While such a structure provides an improved mechanical bond between the insulating ceramic and the substrate, it has several disadvantages which limit its commercial application. For example, a prefabricated honeycomb structure is useful for flat surfaces, but can not conveniently be bonded to a curved surface, such as an airfoil surface. A brazed bond between the reinforcing structure and the substrate creates a temperature limitation due to the low melting temperature of the braze material. A weld bond between the reinforcing structure and the substrate involves the introduction of undesirable heat into the material. Welding and brazing onto a concave or convex surface is very difficult to accomplish on a production basis. Finally, the thermal conductivity of the metal reinforcing structure itself decreases the overall insulating effectiveness of the thermal barrier coating layer.
U.S. Pat. Nos. 5,419,971 and 6,074,706, respectively dated May 30, 1995 and Jun. 13, 2000, describe a thermal barrier coating system having an interfacial layer with a pattern of small, shallow three-dimensional features such as grooves or dimples each having depths and widths of about 12.7 to about 25.4 micrometers. A thermal barrier coating is deposited into the small, shallow features and over the interfacial layer. These systems recognize and accept that cracks form between the interfacial layer and thermal barrier coating, and seek to arrest crack propagation by placing obstacles in the form of the small, shallow features into the crack path. While these systems may or may not achieve their intended result of merely delaying crack propagation, they do not and cannot localize and segregate the cracks to prevent them from combining with other cracks within the system.